In a gas turbine aircraft engine air enters at the engine inlet and flows from there into the compressor. Compressed air flows to the combustor where it is mixed with injected fuel and the fuel-air mixture is ignited. The hot combustion gases flow through the turbine. The turbine extracts energy from the hot gases, converting it to power to drive the compressor and any mechanical load connected to the drive. These hot gases produce temperature differentials that can cause plastic deformation of the turbine casing if the latter is not properly shielded.
The turbine consists of a plurality of stages. Each stage is comprised of a rotating multi-bladed rotor and a nonrotating multi-vane stator. The blades of the rotor are circumferentially distributed on a disk for rotation therewith about the disk axis. The stator is formed by a plurality of nozzle segments which are butted end to end to form a complete ring. Each nozzle segment comprises a plurality of generally radially disposed vanes supported between inner and outer platforms. Each vane and blade comprise an airfoil section.
The abutting outer platforms of the nozzle segments and the abutting outer platforms of the rotor blades collectively define a radially inwardly facing wall of an annular gas flow passageway through the turbine, while the abutting inner platforms of the nozzle segments and the abutting inner platforms of the rotor blades collectively define a radially outwardly facing wall of the annular gas flow passageway. The airfoils of the rotor blades and nozzle guide vanes extend radially into the passageway to interact aerodynamically with the gas flow therethrough.
During operation of the gas turbine engine, it is desirable to minimize thermally induced plastic deformation of the outer casing. This can be accomplished by isolating the outer casing from the heat produced by the hot gases flowing through the turbine.
One source of hot gas leakage into the space between the outer casing and the annular passageway is the interface between the turbine shroud and the hardware which supports the aft edge of the shroud. For example, in U.S. Pat. No. 3,393,894 to Redsell, the turbine shrouds are trapped between the casing and the nozzle. Alternatively, as shown in U.S. Pat. No. 3,542,483 to Gagliardi, the outer shrouds and blade rings can all be hung from the casing with the outer shrouds and blade rings stacked axially in abutting sequence. In accordance with such prior art arrangements, a clearance fit in the axial direction is required to accommodate mechanical stack-up, differential expansion and assembly of the components. Such an axial clearance is susceptible to hot gases leaking therethrough.
One technique for blocking gas flow through an axial clearance, disclosed in U.S. Pat. No. 4,573,866 to Sandy, Jr. et al., is to insert a spring bellows seal in an annular groove between the casing and the nozzle. The seal forms a barrier to gas which has flowed through an axial clearance between the nozzle and the tip shroud support ring.
In the case where a curved circumferential surface of a turbine shroud bears on a curved circumferential surface of an outer platform of a nozzle, as taught in U.S. Pat. No. 3,056,583 to Varadi et al. and U.S. Pat. No. 4,537,024 to Grosjean, a different problem arises. Due to differential expansion during operation, the nozzle outer platform becomes hotter and expands more than the turbine shroud, causing the radius of curvature of the nozzle outer platform to become greater than the radius of curvature of the turbine shroud. The result is a gap between the overlapping turbine shroud and nozzle outer platform, through which hot gases leak from the annular gas flow passageway. These hot gases can produce an undesirable increase in the temperature of the outer casing.